The present invention generally relates to satellite orbital propulsion and satellite power generation systems, and more particularly to a method and apparatus of employing a conducting tether to control satellite system state, such as to produce an electrodynamic propulsion force to change the orbit of a satellite through interaction between an electric current in the tether and an external magnetic field, or to generate electric power onboard the satellite.
Space tethers have attracted a lot of attention in the past 40 years. Many researchers have contributed to the theory of tether behavior in orbit. The theory has been applied and proved in a number of space flights involving tethers attached to spacecraft.
In 1966, Gemini 11 and 12 manned spacecraft were attached with a tether to a rocket stage and demonstrated libration and rotation modes of tethered motion.
In 1992, TSS-1, the Shuttle-based Tethered Satellite System, including a 550-kg satellite and a 20-km electrically conductive tether, was partially deployed from the Shuttle orbiting at an altitude of 296 km. The measurement of the voltage-current profiles shed new light on electric behavior of conducting tethers in orbit.
In 1993, SEDS-I, the Small Expendable Deployment System, including a 26-kg mini-satellite on a 20-km non-conductive tether, was successfully deployed downward from a Delta rocket second stage. SEDS-I was a flight experiment to test the deployment of a long tether by means of a light and simple deployment mechanism and the deorbit and reentry of the mini-satellite after the release of the tether from the Delta stage. The 20-km non-conductive tether was the longest structure ever deployed in orbit.
Also in 1993, PMG, the Plasma Motor Generator, including a 500-m-long electrodynamic tether, was deployed from the Delta second stage with the primary goal of testing power generation and thrust by means of an electrodynamic tether. This mission was the first example of a propulsion system for space transportation that did not utilize any propellant, but rather achieved propulsion by converting orbital energy into electrical energy (deorbit) and electrical energy into orbital energy (orbit boosting).
In 1994, SEDS-II, the Small Expendable Deployment System (second flight), with the same equipment as SEDS-I, was utilized for a longer and more ambitious mission. SEDS-II was stabilized along the local vertical at the end of deployment and kept attached to the Delta stage to study acceleration environment. During the extended mission phase, the deployed SEDS-II was used to study the survivability of a thin tether to micrometeoroid impacts. During the extended mission phase, SEDS-II also provided important data on the micrometeoroid risk as the tether was cut at the 7-km point three days after the completion of the one-day primary mission.
In 1996, TSS-1R, a reflight of TSS-1 was attempted. The mission was terminated before due time by an electrical arc that severed the tether just before the end of deployment. Nevertheless, it was an important mission for tethered satellites, because it showed that the electrodynamic tethers were more efficient at collecting electrons than most theories and models predicted, providing valuable data on electric performance of electrodynamic tether systems.
In 1996, TiPS, the Tether Physics and Survivability Experiment, including a 4-km-long passive tethered system for the investigation of the long-term survivability of tethers in the space environment, was successfully started. This system proved that a sufficiently fat tether can survive for a very long time the harsh space environment, and also provided valuable data on the long-term passive internal damping of tether librations.
In 1999, the Advanced Tether Experiment (ATEx) began deployment in orbit. About 18 minutes into deployment, at a deployed length of only 22 meters, the tether went slack, bent, and triggered several tether departure angle optical sensors. This led to the tether experiment being automatically ejected, to protect the host vehicle. The slackness occurred just after sunrise and may have resulted from a thermal transient on the thin polyethylene tape tether.
In 2002, ProSEDS, the Propulsive Small Expendable Deployer System, is scheduled to deploy 10 km of Dyneema tether followed by 5 km of uninsulated wire from a Delta-II stage to test the electrodynamic propulsion capabilities of the tether. xe2x80x9cTethers in Space Handbook,xe2x80x9d First Edition, NASA Office of Space Flight, NASA Headquarters, Washington, D.C., 1986, edited by P. A. Penzo and P. W. Ammann, provides summaries of various applications and features of space tethers, including methods to change orbital elements with electrodynamic tether propulsion and methods to control the attitude dynamics of such tethers. The basic concept is to vary the electric current in the tether based on the estimate of the tether state obtained from measurements of certain tether system parameters.
The following patents cover certain details of electrodynamic tether usage.
U.S. Pat. No. 6,116,544, entitled xe2x80x9cElectrodynamic Tether and Method of Use,xe2x80x9d issued Sep. 12, 2000, to Forward et al., describes electrodynamic tethers for deorbiting out-of-service satellites.
U.S. Pat. No. 6,260,807, entitled xe2x80x9cFailure Resistant Multiline Tether,xe2x80x9d issued Jul. 17, 2001, to Hoyt et al., discusses various multistrand tethers to improve strength and stability.
U.S. Pat. No. 4,923,151, entitled xe2x80x9cTether Power Generator for Earth Orbiting Satellites,xe2x80x9d issued Mar. 1, 1988 to Roberts et al., discloses use of an electrodynamic tether as a power generator for earth orbiting satellites.
U.S. Pat. No. 4,824,051, entitled xe2x80x9cOrbital System Including a Tethered Satellite,xe2x80x9d issued Jan. 12, 1987 to Engelking, discloses use of an electrodynamic tether attached to a satellite to compensate for the air drag and the orbit degradation.
U.S. Pat. No. 3,868,072, entitled xe2x80x9cOrbital Engine,xe2x80x9d issued Feb. 25, 1975, to Fogarty, discloses a tether to rotate/revolve one mass about the other and provide energy.
U.S. Pat. No. 3,582,016, entitled xe2x80x9cSatellite Attitude Control Mechanism and Method,xe2x80x9d issued Jun. 1, 1971, to Sherman, discloses a study about transverse waves and rotational dynamics. The ""016 Patent does not disclose electrodynamics or use of magnetic fields.
Most of the early estimates of performance of electrodynamic tethers were based on static stability considerations, where non-stationary processes were ignored. In recent years, however, more attention has been given to dynamic stability considerations, where complex non-stationary dynamic response to real perturbations is taken into account.
V. V. Beletsky and E. M. Levin in xe2x80x9cDynamics of Space Tether Systems,xe2x80x9d Advances in the Astronautical Sciences, v. 83, AAS, 1993, described many modes of inherent instabilities of electrodynamic tethers that are observed even in equatorial circular orbits, and even when dynamic models neglect magnetic field variations along the orbit. The Beletsky and Levin reference points out that it would be virtually impossible to operate electrodynamic tether systems anywhere close to the boundaries of static stability, because of a very strong, uncontrollable or hardly controllable dynamic instability in these regions. It has been shown in this study that realistic expectations for safe electric current levels must be typically lowered by an order of magnitude compared to static levels because of dynamic instabilities.
More evidence of rigid dynamic instability constraints was accumulated, as more detailed and realistic simulations were performed.
J. Pelaez, E. C. Lorenzini, O. Lopez-Rebollal, and M. Ruiz in xe2x80x9cA new kind of dynamic instability in electrodynamic tethers,xe2x80x9d AAS 00-190, AAS/AIAA Space Flight Meeting Jan. 23-26, 2000, pointed out that instability is inherent in any uncontrolled electrodynamic tether motion and simply cannot be avoided.
R. P. Hoyt and R. L. Forward in xe2x80x9cThe Terminator Tether: Autonomous Deorbit of LEO Spacecraft for Space Debris Mitigation,xe2x80x9d AIAA 00-0329, 38th Aerospace Sciences Meeting and Exhibit, Jan. 10-13, 2000, Reno, Nev., reported that active control had to be applied to cope with dynamic instabilities. The results of their detailed dynamics simulation results showed actual performance levels much more conservative than described in the Forward et al. U.S. Pat. No. 6,116,544 based on static stability considerations. This was attributed to the dynamic stability constraints.
J. Corsi and L. less in xe2x80x9cStability and Control of Electrodynamic Tethers for De-orbiting Applications,xe2x80x9d LAF-00-S.6.06, 51st International Astronautical Congress, Oct. 2-6, 2000, Rio de Janeiro, Brazil, showed that realistically for deorbiting with an electrodynamic tether, the electric current in the tether must be periodically switched off to prevent libration buildup and rotation onset, thus substantially decreasing deorbiting efficiency of the electrodynamic tether.
R. P. Hoyt, G. Heinen, and R. L. Forward in their presentation on the xe2x80x9cProgress on Development of the Terminator Tether,xe2x80x9d NASA JPL/MSFC/UAH 12th Annual Advanced Space Propulsion Workshop, University of Alabama in Huntsville, Huntsville, Ala., Apr. 3-5, 2001, took this idea further, indicating that it may be sufficient to reduce the current by 10% when the tether swings in-plane backwards to prevent instability. No control of out-of-plane and skiprope modes was attempted.
It is important to prevent not only any strong instability, but also weak instabilities that will impact long-term electrodynamic tether missions. In other words, electrodynamic tethers in most missions must be stabilized on a long-term basis.
One of the most commonly suggested control strategies is to vary Ampere forces distributed along the tether by varying the electric current in the tether. G. Colombo, M. Grossi, M. Dobrowolny, and D. Arnold in xe2x80x9cInvestigation of Electrodynamic Stabilization and Control of Long Orbiting Tethers,xe2x80x9d March 1981, SAO, Contract NAS8-33691, discussed some strategies of how to use electric current modulation to stabilize an electrodynamic tether in a near-vertical position.
E. Netzer and T. R. Kane in xe2x80x9cUses of ElectroDynamic Forces in Tethered Satellite Systems,xe2x80x9d 1994, showed how to augment the electric current control with thruster pulses to keep the electrodynamic tether close to the local vertical during orbit transfers.
In order to control tether dynamics, the tether libration and oscillation state must be properly estimated throughout the mission.
A combination of satellite attitude angles, attitude rates, and gyro data, along with Earth sensor and magnetometer measurements was used to estimate the tether state in the TSS flights, as reported by H. Biglari and Z. J. Galaboff in xe2x80x9cAn Extended Kalman Filter for Observing the Skiprope Phenomenon of the Tethered Satellite System,xe2x80x9d 4th International Conference on Tethers In Space, Washington, D.C., Apr. 10-14, 1995.
E. Netzer and T. R. Kane in xe2x80x9cEstimation and Control of Tethered Satellite Systems,xe2x80x9d 1994, suggested measuring angles between the local vertical and the tether at the attachment points to estimate the tether state, including rotation and bending.
In 1995, Gullahorn et al, in xe2x80x9cLong Period Tension Variations in TSS-1 and SEDS-2,xe2x80x9d in the proceedings of the Fourth International Conference on Tethers in Space, point out that xe2x80x9cobservations of tension in a tethered system could provide a measure of librational amplitude and phase.xe2x80x9d The analysis of flight data they present shows that in-plane libration is typically far easier to detect than out-of-plane libration using this technique. This is expected because in-plane libration induces large Coriolis effects while out-of-plane libration does not. On SEDS-2, a consistent pattern of a +/xe2x88x928% tension variation with a 53 minute period suggests a 4 degree amplitude in-plane libration after the end of deployment.
While these estimation and control approaches for tethered satellite systems were suited to work in their respective particular cases, there is a need for more general and more practical estimation and control approaches that will work not only in the vicinity of the local vertical, but also over a wider range of dynamic conditions encountered by electrodynamic tethers. In addition, there is a need for estimation and control approaches for tethered satellite systems that employ more simple and readily available measurements to estimate the tether state.
One aspect of the present invention provides a tether system for space applications including at least one electron collector, at least one electron emitter, a power system, and at least one electrodynamic tether electrically coupled to the power system. The electrodynamic tether conducts electrical current between at least one electron collector and at least one electron emitter. Electric current and voltage are measured. Selected tether dynamics and electrodynamics based on the measured electric current and voltage are ascertained. The electric current and/or current distribution is adjusted to thereby control at least one tether system state based on the ascertained selected tether dynamics and electrodynamics.